Combined aileron and landing flap



R. E. CRANDALL COMBINED AILERON AND LANDING` FLAP June 29, 1954 5 sheets-Sheet 1 Filed oct. 50, 41948 NNN June 29, 1954 R. E. CRANDALL 2,682,381

COMBINED AILERON AND LANDING FLAP Filed Oct. 30, 1948 5 Sheets-Sheet 2 N .s M5 wien/r irre/Mey June 29, 1954 R, E, @RANDALL 2,682,381

COMBINED AILERON AND LANDING FLAP Filed Oct. 50, 1948 5 Sheets-Sheet 3 AcuAraa Umz- 67 IN VEN TOR. ,Qa/144m 69mm@ June 29, 1954 R E CRANDALL 2,682,381

COMBINED AILERON AND LANDING FLAP Filed OCT.. 50, 1948 5 Shee'CS-Sheet 5 IN V EN TOR. 0A/4m mA/ma BY y Patented June 29, 1954 COMBINED AILERN AND LANDING FLAP Ronald E. Crandall, Los Angeles, Calif., assignor to Northrop Aircraft, lne., Hawthorne, Calif., a corporation of California,

Application October 30, 1948, Serial No. 57,518

4 Claims.

This invention relates to airplane control systems, and more particularly to a means for combining two functions, operated by two independent controls, into a single flight control surface of an airfoil.

In high speed airplanes with small wing areas and high wing loadings, it becomes increasingly more difficult to attain low landing speeds, for the main reason that landing iiap area cannot be made large enough to contribute the desired extra lift. The leading edge of the landing flap cannot be moved forward to widen the flap without cutting through the rear wing spar which must reach to the fuselage or body section. The landing flap cannot be lengthened in the outboard direc- Y tion beyond a certain limit without sacrificing necessary aileron surface area.

It is, therefore, an object of the present invention to provide a method for increasing the lift of an airplane in flight without compromising on other desirable features.

Another factor of importance concerning present-day aircraft controls systems is the problem of handling very large or very heavily loaded control surfaces by pilot control stick forces small enough to be possible or practical under conventional piloting procedures. Control surface- Vmounted servo tabs have been extensively used to assist the pilot in overcoming these high control forces, but more recent developments have been along the line of power-operated systems employed in conjunction with direct control cable connections to surfaces having no servo tabs.

Another method recently adopted, notably in the United States Air Force heavy bomber designated as the B-35, manufactured by the assignee of the present invention, is the use of full-power flight controls wherein the pilots control elements merely introduce signals into power-driven actuators connected to the movable surfaces. This full-power method is ideally suited for an aircraft control system because of its irreversibility and overall simplicity as compared to a powerboost system.

Recognizing these advantageous features of the full-power control system method, it is a further object of this invention to apply the principles of duplication of functions performed by a single airfoil to such a full-power system, or

systems, and thereby to provide a control system V to independently deflect an attitude control surface, which is normally operated by one of two alternate and smoothly substitutable full-power control systems, without affecting the range of movement, operating characteristics, or the operation of the normal systems in any way after the newly deflected position is reached, or at any time while the separately deecting action is in motion. ln this way, for example, full-span landing flaps can be achieved by combining aileron drooping with flap panel lowering, and a detailed description of this embodiment is contained in this specification.

The specific apparatus which embodys the present method of control to be described herein comprises a full-power aileron control system normally controlled by either of two power mechanisms, with a mechanism added to move the aileron independently of normal roll control by energizing that power system which happens to be engaged to drive the aileron, this added movement being in a substantially linear relation to movement of a separate landing flap. Other surfaces or systems can be combined equally as well, using the same method. Therefore, it is to be understood that this invention is not limited in any way to the apparatus of the present application.

Reference is made to the following drawings, in which:

Figure i is a perspective view showing the inboard end of an aileron control mechanism connected to operate one aileron of an airplane.

Figure la is a perspective view in continuance of Figure 1, showing the outboard end of the same control mechanism.

Figure 2 is a perspective diagrammatic view showing the internal construction of a control switch box for operating an electrical aileron actuator.

Figure 3 is a block diagram showing the components of an electric aileron actuator unit.

Figure 4 is a schematic electrical diagram showing the wiring of an electrical control system for an airplane attitude control surface.

Figure 5 is a schematic diagram showing the basic components involved in shifting from electrical to hydraulic full-power operation of an -airplane attitude control surface, or Vice-Versa.

Figure 8 is a perspective diagrammatic View showing the output end of an airplane landing iiap control system.

Figure 7 is a perspective View showing on a larger scale than in Figure 1 the connection of the flap control system of Figure 6 to the aileron control system.

Referring first to Figure l, two aileron cables l,l are conventionally connected between a pilots control stick 2 and a quadrant assembly 3 located near the aileron to be operated, so that lateral movement of the control stick 2 will produce rotational movement of the quadrant assembly 3. Before reaching the quadrant assembly 3, the cables I,I pass over pulleys 4,4 of a cable tension regulator 5 designed to maintain a substantially constant tension with varying temperature conditions. The quadrant assembly 3, which is pivoted to rotate about a fixed vertical axis 6, has two arms 'Iy and iiA attached thereon which project horizontally from the quadrant axis 5 approximately 90D apart. A horizontal aileron control rod 9 is pin-connected. to the end of the quadrant control arm 'I and to a control rod arm Ill of a bell crank assembly II. Bell crank assembly II also includes a valve control crank I9 formed integrally with a crank leg 3l) which is positioned just below the control rod arm IB. These crank assembly components it, i9 and Sil are pivotally mounted on an axis pin I3 which is xed to the airplane structure and is parallel tothe quadrant assembly axis 6. Control rod arm Ill is connected to valve control crank I3 in a manner to be described later. To the end of the quadrant spring arm 8 is pin-connected arhorizontal spring-loaded centeringr rod I2 which isirotatably attached to axis pin I3.

When the quadrant assembly 3 is in the neutral position, the aileron control rod 9 and its quadrant control arm 'I form an angle of about 90, while-the centering rod I2 and its quadrant spring arm 3 form a straight line. Therefore, when the control stick 24 is moved laterally from the neutral position, the aileron control rod 9 will rotate the control rod arm IS of the bell crank assembly i I about its axis pin i3, and the centering rod l2 will be stretched and exerting a turning moment on the quadrant assembly 3 tending to return the quadrant assembly 3 and the entire system to neutral. The centering rod I2 produces a synthetic feel to the control stick, and the farther the stick ismoved from the neutral aileron position, the larger will be the restoring force.

Also connected to the bell crank axis pin I3 is a piston rod i4 oi' a hydraulic actuating cylinder I5.4 This actuating cylinder i5 has a servo control valve IS mounted on the cylinder housing. The control valve IE is provided with a valve shaft` Ilwhich extendsfroin the control valve parallel to and at the same end of the cylinder I3 as the piston rod I4. A valve operating rod IE connected between the valve shaft I'l and a valve control crank I- of the bell crank assembly Ii. rThe valve control crank i9 and the control rod arm I of the bell crank assembly II are connected together in such a manner that the vertex angle between them can be changed by an independent outside action. This connection and its operation will be described in detail later.

The closed end of the actuating cylinder I5 opposite the piston rod I4 is supported by and pin-connected toa cylinder support arm 2d or an inboard lever assembly 2l which rotates about a xed vertical axis. An upper arm 22 of this lever assembly 2l carries a link 23 which is connected to an aileron tting 24 attached to the aileron spar 25, so that longitudinal translation of the link 23 will cause aileron deflection about the aileron yhinge line 26.

Hydraulic pressure 2l and return 21a lines are supplied from the airplanes hydraulic system to the control valve I6 and fluid under pressure is directed' by the valve to the proper side of the enclosed piston (not shown) attached to the piston rod I4 of the actuating cylinder I5, when the valve shaft 'II is displaced from its neutral position, to move the cylinder I5 relative to the piston rod I4. By proper side of the piston is meant that side where the applied pressure will move the cylinder housing in the same direction that the valve shaft IT was moved. Since the piston rod I4 is fixed to the rigid bell crank axis pin I3, and since the servo control valve I6 is mounted on the cylinder housing I5, it may be readily understood that the cylinder housing I5 will move only as far as the valve shaft was originally displaced, since movement of the control valve IG with the cylinder housing I5 will return the valve shaft I'I to neutral with respect to the valve. The system is particularly designed sol that if the surface load at some deflected position exceeds a pre-determined danger value, the hydraulic fluid flow will be stalled and pushed back into the high pressure line until a smaller deflection is reached where the air load is lowered andthus kept below a safe maximum. This full-power hydraulic control system is the subject of an abandoned application, Serial Number 23,567, filed April 27, 1948, and owned by the assignee of the instant application.

One hydraulic actuator is used per aileron, and a similar system cable-controlled by the same control stick is installed to actuate the aileron surface on the other side of the airplane, as indicated by cables Ia, in. Figure l. The aileron system thus far described is a complete full-power hydraulic control system capable of satisfactorily controlling the attitude of two roll control surfaces of an airplane. It will now be shown how a second or standby system is combined with the iirst to be operated from the same control stick, but to operate the aileron by electrical means.

Connected to the crank leg 3B, previously described as being formed integrally with the valve controlcrank I9', is a-switch control rod 3| which is in turn connected to a control lever 32 of a switch box 33. In this manner, the control lever 32 will be moved in exact accordance with the valve control crank I9 which, of course, is controlled by the pilots control stick 2 through the quadrant assembly 3, aileron control rod 9, and controlrod arm I0. The switch box 33 has a follow-up lever 34 which is connected by a follow-upfrod 35 to one end of a follow-up crank 3G pivoted on the same axis pin I3 as the bell crank assembly II. The other end of the follow-up crank 36is connected by a push-pull rod 37 to an eye fitting. 38 .installed on the actuating cylinder housing I5. Thus, the follow-up lever 34 is made to assume a denite position corresponding to the aileron position, since the cylinder housing I5 is connected with the aileron 25 as described above.

The construction of the switch box 331s shown in Figure 2, which shows the box in an upright position with the two levers 32 and 34 projecting from the top. The control lever 32 to which the switch control rod 3l is connected at the top is rotatably attached at a point approximately one-third of its length down from the top to a sub-lever pin 40 extended from a sublever 4I 'which has the same motion as the follow-up lever 34, since the two are rigidly connected together at their lower ends by a connection shaft 42, which forms the axis of rotation for the follow-up lever 34. The lower end of the control lever 32 connects by a link pin 43 to a link bar 44v in turn pin-connected to a crank arm 45 which is welded to a control shaft 48. The control shaft 46 is mounted in stationary end bearings 41, 41 so that it is free to rotate about its longitudinal axis. Along the length of this control shaft are welded various links 48 which are pin-connected to other links 45 whose purpose it is to effect the operation of five micro switches 5| through 55 mounted securely in adjustable brackets (not shown) in the switch box 33. These links are arranged as follows: an idler shaft 56 is positioned in the switch box 33 so it is parallel to the control shaft 46 and located directly under the ends of the micro switch actuators 51. Eight vertical links 58 are separately pivoted on the idler shaft 56 and extend upwardly so that their top ends are immediately adjacent to the switch actuators 51 in a position to move the actuators, one such link for the first switch 5|, one link for the second switch 52, and two links each for the third 53, fourth 54, and fifth 55 switches. Connected to the top of each vertical link 58 is a connecting link 49 which is pinned to one of the welded links 48 on the control shaft 46.

When assembled, the connecting links 49 and the welded links 48 occupy the space between the tops of the vertical links 58 and the control shaft 46, as shown in Figure 2, but do not form a straight line connection. For the rst switch 5|, the connecting link and the welded link slantI up toward each other, and for the second switch 52, the connecting link and the welded link slant down toward each other. For the three remaining switches, which have two sets of associated links instead of only one, one set for each of these switches is arranged identically with the set for the first switch 5|, and the second set for each remaining switch is arranged identically with the set for the second switch 52.

When the mechanism is assembled in this manner, when the top of the control shaft 46 is turned toward the micro switches, the rst 5|, third 53, fourth 54, and fifth 55 switches will be actuated by the link sets which approach a straight line and thereby force the respective switch actuators inwardly to close the circuit. When the top of the control shaft 46 is turned away from the micro switches, the second 52, third 53, fourth 54, and fth 55 switches will be actuated by the remaining link sets which took nc part in the previous operation.

The position of each switch on its support bracket is carefully adjusted and then locked so that as a whole they are actuated progressively when the control shaft 46 is turned in either direction from neutral, either the first 5| or second switch 52 being closed first, depending on the direction of rotation, and then the third 53, fourth 54, and fifth 55 switches in the order named. As. will be explained further later, the rst and second switches are direction and low speed switches, i. e., revolve the controlled motor at low speed to apply either up or down aileron, and the remaining three switches are speed control switches which serve to increase the speed of aileron movement with increasing angle of rotation of the control shaft 46 from the neutral position.

The control lever 32 of the switch box may pivot around either of the two points, i. e., the sublever pin 40 or the upper connection of lever 32 with the switch control rod 3|, and this action makes the follow-up action possible. When the` switch control rod 3| moves the control lever 32 toward the bell crank assembly for example, as when the control stick 2 is moved, the control lever 32 will pivot about the sub-lever pin 40, since the aileron at this time is still stationary and, as described, is directly connected to the sub-lever 4| through the cylinder housing |5, the follow-up crank 36, the follow-up rod 35, followup lever 34, and connection shaft 42. This action of the control lever 32 will move the lower end of the crank arm 45 to the left in Figure 2, causing a clockwise rotation of the control shaft 46, and thus actuating one or more micro switches. As a result, the aileron is in motion, and it is connected to push the follow-up rod 35 in the same direction which the switch control rod 3| was pushed; in this instance, to the right, in Figure 2. As this occurs, the sub-lever 4| will act on the control lever 32 to rotate it about the end of the switch control rod 3|, since this upper point is now being held solidly by the control stick. When the control lever 32 pivots in this manner, its lower end will be moved back to the right, in Figure 2, turning the control shaft 46 counterclockwise to its original position, where the micro switches are opened in the reverse order of their closing, and the aileron is stopped at some Ideflected position. If the control stick is now moved back to neutral, the reverse action will take place, and the opposite directional switch will be closed to return the aileron to neutral, at which time the switches will be off again. Thus, the aileron surface is made to respond to stick movement and assume definite positions corresponding to various stick displacements. The three speed control switches 53, 54, and 55 function to make the speed of aileron movement closely simulate the speed of control stick movement.

As shown in Figure l, a lower arm 6|] of the inboard lever assembly 2| is pin-connected to a tube assembly 6| which is located in the wing forward of the aileron 25, and extends from one end of the aileron to the other end, as further shown in Figure la. At this other end, tube assembly 6| is pin-connected to a lower crank 62 of an outboard lever assembly 63, mounted to rotate about a fixed axis similar to the inboard lever assembly 2|. An upper crank 64 of this outboard lever assembly 63 connects by a link 23a to another aileron fitting 24a to form a second point of application of control force to the aileron spar 25.

An actuator support arm 65 of the outboard lever assembly 63 is pin-connected normally at a right angle to a drive screw 66 of an electrical actuator unit 61. The actuator unit 61 is connected to a load limiter 68 which is mounted rigidly to the aircraft structure. Extension or retraction of the drive screw 66 by the actuator unit 61 will, therefore, operate the aileron 25 through the same links 23 and 23a as when the aileron is operated by the hydraulic actuating cylinder l5, since both the drive screw 66 and the actuating cylinder |5 are positively connected to both lever assemblies 2| and 63 by means of tube assembly 6|. The actuator unit 61 is normally controlled by the switches in the switch box 33, and its construction and operation will now be described.

Electrical actuator unit Figure 3 diagrammatically shows the internal relationships of the components of the electrical actuator unit. One end of the actuator unit 61 is pivotally attached to a load limiter 66. Inside the actuator unit is a compound wound D. C. reversible electric motor 69 which provides motive power to drive the aileron 25. The motor 69 has a motor shaft 10 which is' connected to drive a two-speed gear train 1'|` through a friction clutch 12. The gear train 1| is provided with a multiple disk brake 13. Both the motor clutch 12 and the gear train brake 13 are controlled and operated simultaneously by a solenoid A whose arm 14 is acted on by a tension spring 15. When this solenoid is de-energized, the spring 15 will engage the brake 13' to apply braking force to the gear train 1|, and at the same time, release the clutch 12 from the motor S9. When solenoid A is energized, its arm 14 will be magnetically pulled against the pull of the spring 15 so that the gear train brake 13 will be released and the motor clutch 12 will be engaged in driving position.

The gear train 1| drives an output shaft 15 to which is ixed a drive screw nut 11. Drive screw nut 11 turns on drive screw 66 which is made to extend or retract with the rotation'of drive screw nut 11, and thus drive the aileron through tube assembly 3| and the lever assemblies 2| and 63. This drive screw and nut are of the ball bearing type and are reversible; that is, the nut portion can be revolved by axial loads on the screw portion, such as caused by air loads on the aileron or by movement of the hydraulic actuating cylinder |5.

The two-speed gear train is governed by solenoid B and solenoid C, both spring--loaded in the 11i de-energized position. When neither of these two solenoids is energized, the gear train is disconnected within itself so that the drive screw nut 11 may free wheel in response to a push or pull on the drive screw S6 without turning the I entire gear train. When only solenoid C is energized, the gear train is connected to rotate the drive screw nut 11 at low speed. When only solenoid B is energized, the gear train ratio is changed to rotate the drive screw nut at high speed. i

In electrical operation, when no micro switches in the switch box 33 are closed, solenoid C only is energized, the gear train brake 13 is set, and the aileron is immovable, as shownin Figure 4. 'When a directional switch 5| or 52 is closed, solenoid A and the motor circuit are simultaneously energized (with solenoid C still energized), which releases the gear train brake 13 and starts the motor 59 in the proper direction to drive the aileron at low speed, designed to be approximately 5 oi angular movement per second. If the other directional switch is closed instead, the leads to the brushes on the motor 69 are reversed to cause the motor to run in the opposite direction, as shown in Figure 4. The electric motor 69 has nve wire leads. Wire M5 always connects to ground. For reversing the direction of motor rotation, the connections of the armature wires M2 and M3 are reversed. Thus, for operation in one direction, as controlled by the iirst switch i, wires Mi and M2 are connected together and wires M3 and M4 are connected to the positive side of the D. C. line'. For operation in the opposite direction, as controlled by the second switch 52, wires MI and M3 are connected' together and wires M2 and M4 are connected to the positive side of the D. C. line. The proper connections are made by switch relays 19 and motor relays 80 which are operated by the directional switches 5| and 52 in the control switch box 33'.

When the iirst speed control' switch 53 is closed' in addition to the directional switch, solenoid B is additionally energizedan'd solenoid'C is` de- 8 energized, which action' shifts the gear ratio of the system to move the aileron faster, at approximately 13 angular movement per second, for example, while the motor continues to revolve at the same initial speed.

When the second speed control switch 54 is closed, an external resistance R| is added to the shunt held 8| of the motor, causing the motor to turn faster, using the same gear ratio as b'efore, as shown in Figure 4. Preferred aileron surface velocity is now approximately 21 per second. u

When the last speed control switch 55 is closed, a greater additional external resistance R2 is added in series with the shunt field 8|, causing a still higher motor speed. A preferred maximum aileron velocity of approximately 30D per second is now attained. Y

It will be noted that when novpower at all is supplied to the switch box 33 to control the actuator circuit, the aileron surface is not locked at any position, but may be moved up or down, causing the actuator drive screw nut 11 to free wheel separate from the gear train 1|. This full-power electrical control and follow-up system is the subject of a co-pending application,

Serial Number 65,806, iiled December 17, 1948,`

and owned by the assignee of the instant application.

Up' to this point, a complete hydraulic fullpower system and a complete electrical fullpower system have been described, either of which, when energized byits respective power supply, is connected to operate the same aileron when controlled by the same control stick, and in a'manner which gives substantially the same operating characteristics as a direct manual control system, if the non-controlling system is suitably bypassed. y

Figure 5 shows diagrammatically how shifting of control from one system to the other can be accomplished. A two-position manual control switch 85 is provided at the flight station. The closed position of this switch selects the electrical control system and the open position selects the hydraulic control system. In the open position, no energy is supplied to the switch box 33 therelore, it cannot be operated, but the actuator drive mechanism will free wheel with any motion of the aileron surface, as described before. At the same time, the hydraulic system is in normal operation, and can control the aileron as usual. In the closed position of they manual control switch 55', electrical energy is supplied to both the switch box control circuit and the motor operating circuit, so that the entire electrical system is cut into operation. Simultaneously with the switch closing, two things happen in the hydraulic system; (1) A solenoid bypass valve 86 is opened, mutually connecting each side of the piston 81 in the hydraulic actuating cylinder |5. (2) A solenoid shut-off valve 88 is closed, shutting off the hydraulic pressure supply line 21 at the servo control valve I6.` This action allows free wheeling action of the piston 81 without causing a hydraulic lock, and prevents the pressureline 521 from being continuously bled into the return line 21a. In actual practice, it is preferred that the above two valving operations be controlled by a single solenoid mounted on the control valve i6, since' in the present apparatus the control valve I6v and actuating cylinder |5' are enclosed in a single housing.

In addition tothe manual? control switch 85;

automatic' emergency switching means is pro-l vided to shift from hydraulic to electrical operation if for any reason the fluid pressure drops below operating level. This is accomplished by including a pressure-operated switch 89 in the hydraulic pressure line 21, this switch being in A parallel with the manual control switch 85, and

adjusted to close when the pressure failure occurs.

It will thus be seen that both systems are in position at all times to be energized and assume control smoothly without affecting overall operating characteristics or producing a jerk in surface position orn pilots controls when the shift is made. This combination of two fully powered control systems alternately connectable to operate an airplane attitude control surface from the same single signal-producing input is the subject of a co-pending application, Serial Number 42,265, filed August 3, 1948, now Patent No. 2,638,736.

Having now shown the use of this latter-menu tioned combination, the present invention consists of a mechanism to cause that same attitude control surface to deflect in relation to the movement of an independently controlled device, utilizing as a power means for this deflection whichever alternate `full-power system is in normal control at the time of such movement. The independently controlled device in the specific apparatus shown herein is an airplane landing flap system, the output end of which is shown in Figure 6. A flap drive shaft 90 rotated by the flap drive motor M connects through a gear box 9| to a flap jackscrew 92 which rotates in a nut unit 99 attached to swivel on the framework of a landing flap panel 94. The nap panel 94, being hinged at its leading edge 95, delects downwardly at its trailing edge 96 when the jackscrew 92 screws the nut unit 93 away from the ap drive shaft 90. The nap drive motor M may be operated by any suitable means under the control of the airplane pilot.

An extension 91 of the flap drive shaft 90 leads from the gear box 9| to drive a cable drum 9S by means of worm gears 99. A flap cable wrapped around this cable drum 98 leads to and around a second cable drum |0I located near the aileron quadrant assembly 3, shown in Figure l. The second cable drum IOI drives a universal joint |02 connected to a square tube |03. Thus, when the flap panel is raised or lowered by the turning of the flap drive shaft 90, the square tube |03 will rotate proportionately to the ilap movement.

This square tube |03 is -made in two sections, an inner solid section |04 which is connected to the cable drum |0I, and an outer hollow section |05 which fits over and can slide longitudinally on the inner section |04. The far end of the outer section |05 is connected with a second universal joint |02a to a threaded screw |06, half of which has a right-hand thread and half a left-hand thread. The threaded screw |06 is connected to the bell crank assembly II in the following manner:

The bell crank assembly I| consists of two main parts, as shown in Figure 1, the upper part being the control rod arm I0 pivoted near one end to the bell crank axis pin I3, and the lower part consisting of the valve control crank |9 at one end and the crank leg 30 at the other end, this lower part being pivoted near its center to the bell crank axis pin I3 also. Each bell crank part carries a fork |01 and I01a, respectively, these forks positioned so that when the parts are assembled, a line through the center of the open space of each fork will lie in a horizontal plane parallel to the plane of rotation of each bell crank part and parallel to the sides of the forks. Vertically placed in each fork is a square nut |00 and |99, so installed that each is free to turn about a vertical axis extending from one side of its fork to the other. The nuts |08 and |09 are tapped with opposite-hand threads t0 mate with the threaded screw |06 attached to the square tube |03 is installed through the nuts connected to the forks, the left-hand threaded portion. of the screw turning in one nut, and the right-hand threaded portion of the screw turning in the other nut. A bell crank spring I0 is preferably connected from the control rod arm I9 to the valve control crank I9 to keep the nut connection free of end play.

In this manner, the control rod arm I0 is connected solidly to the valve control crank i9 and crank leg by the threaded screw |06 when the landing flap panel 94| is stationary, During normal aileron movement, the bell crank assembly II acts as a solid member, and as it pivots about its axis pin |9, the square tube |03 can lengthen or shorten without binding because of its telescoping action and the universal joints |02 and I02a at the ends of the tube |03. However, during lowering of the landing flap 94, the square tube |03 will rotate and screw the two nuts |08 and |09 away from each other. Assuming for the moment that the control stick 2 is stationary, the aileron control rod 9 and, therefore, the nut |08 attached to the control rod arm I9 will be held stationary. Therefore, this rotation of the threaded screw |06 will actually elongate the square tube |03 and move the one nut |09 only. When this occurs, rotation of the valve control crank I9 and the crank leg 90 of the bell, crank assembly about the bell crank axis pin I3 will pull the valve shaft I1 out of the control Valve I9 and push the switch control rod 3| toward the switch box 39, which is the proper direction to lower the aileron 25, using whichever power system is in control at this time.

Since the bell crank assembly II is connected directly to the servo control valve I6 through the Valve operating rod I8 only, less frictional resistance to movement of the valve occurs than from the bell crank assembly to the control stick 2. Therefore, practically no force at all is transmitted to the control stick by rotation of the square tube 03. Even if a force of some magnitude were imposed on the control stick through the connecting linkage from this one aileron, an equal and opposite force would also be exerted on the control stick from the aileron on the other side of the airplane, so that no resultant motion is imparted to change the roll control of the ailerons.

The square tube |03, as rotated by the landing flap mechanism, has acted on the same motionproducing members as the aileron control rod 9 would normally act upon, without affecting the position of the aileron control rod 9. A similar extension of the landing flap system on the other side of the airplane works on the opposite aileron mechanism in the same manner, so that as the landing flaps are simultaneously lowered, the ailerons are both lowered also.

It may easily be seen that both normal aileron control and aileron drooping by the flap mechanism can proceed at the same time, when necessary, since operation of one mechanism in no way interferes with the operation of the other. In

fact, full range of aileron control on both sides of the new neutral position can still be obtained even after the ailerons are drooped the maximum amount along with the landing flaps. When the landing flaps are retracted, the ailerons will be power-moved back to their normal neutral position by the reverse action of the bell crank assembly l l from that described above.

With the use of the present method, full-span landing flaps can be achieved while still retaining normal control of other wing-mounted surfaces. While this invention is described as being incorporated on a conventional type airplane with tail surfaces, it also applies to all-wing or other types of aircraft, wherein any such additional surfaces as would be deemed practical can be combined in the same manner at the same time for any purpose. In addition, this invention provides a method for controlling and using a single surface to accomplish landing flap, dive brake, and elevator or roll control functions, since erratically varying hinge moments do not cause variation of the control forces applied by the pilot.` Furthermore, the same method can be used in any other situation where it is desired to couple two independent control sources to the input of a powerdriven device.

Reference has been made herein to the pilot of the airplane as being human. Obviously, the

system as herein described is ideally adapted for control by automatic piloting devices. In fact, the low and uniform control forces required for full-power operation of highly loaded control surfaces, as described herein, make the system readily adaptable for control by automatic pilots,

and the same power units can be used for both human and autopilot control of the surfaces.

From the above description it will be apparent that there is thus provided a device of the character described possessing the particular features of advantage before enumerated as desirable, but which obviously is susceptible of modification in its form, proportions, detail construction and arrangement of parts without departing from the principle involved or sacricing any lof its advantages.

While in order to comply with the statute, the invention has been described in language more or less specific as to structural features, it is to be understood that the invention is not limited to the specific features shown, but that the means and construction herein disclosed comprise the preferred form of several modes of putting the invention into effect, and the invention is, therefore, claimed in any of its forms or modifications within the legitimate and valid scope ofthe appended claims.

What is claimed is:

1. In an aileron and flap control mechanism for an airplane having a pilot controlled aileron actuating mechanism for moving right and left ailerons in opposite directions and flap actuating mechanism for moving right and left flaps in the same direction: a double-armed crank for each aileron, each said -crank being adjustable to change the angle between the arms thereof and one arm of each crank being operatively connected to the right and left aileron actuating mechanism, respectively; aileron control means operatively connected to the other arm of each said crank for rotating said arms in unison to move said ailerons in opposite directions; and means operated by lowering and raising of said iaps connected to each crank for changing the angle between the arms thereof to move said ailerons in the same direction.

2. Apparatus in accordance with claim 1 wherein said latter means comprises a differential mechanism at each of said cranks, a flap followup member, and a linkage having an irreversible driving connection from said follow-up member to said differential mechanism.

3. In an airplane having a pilots roll control member, right and left roll control surfaces, and a lift flap; the combination of nap actuating means operatively connected to said flap to lower and raise said flap; full power servo actuating means operatively connected to each roll control surface and having a control element therefor movable in two directions to energize said power means in either of two directions; a pair of twoarm cranks, one for each said control element, each said crank having one of its arms connected to the corresponding control element and the other of its arms to the pilots roll control member for movement of said control surfaces in opposite directions by movement of the control member; each said crank being adjustable to change the angle betweenv the arms thereof; differential means connected to each crank for changing the angle between the arms thereof; and driving means operatively connected from an output member of said flap actuating means to each of said differential means to change said angles to move the control surfaces in the same direction.

4. Apparatus in accordance with claim 3, wherein said driving means comprises a rotatable driving shaft assembly driven by said flap actuating means, said shaft assembly including two sections connected by a universal joint, and a third section having a longitudinal telescoping connection to one of said nrst two sections, whereby said shaft assembly can have angular misalignments and can longitudinally contract and extend without binding, to permit rotation of said crank carrying differential means.

References Cited in the le of this patent UNITED ,STATES PATENTS Number Name Date 1,199,036 Hodgkinson Sept. 19, 19,1 6 1,300,047 Thomas Apr. 8, ,1919 1,816,787 Moss July 28, 1931 2,183,279 McCarty et al Dec. 12, 1939 2,220,194 Albright Nov. 5, 1940 2,286,150 Mercier June 9, 1942 2,344,594 Bryant Mar. 21, 1944 2,376,731 Stoner May 22, 1945 2,422,035 Noyes Ju'ne 10, 1947 2,445,343 Tyra July 20, 1948 2,479,619 Hilton et al Aug. 23, 1949 FOREIGN PATENTS Number Country Date 598,520 Germany June 1,2, 1934 609,909 Germany Feb. 26, 1935 

